Payload Module Research Articles

Temperature mapping methods for thermoelastic analyses of the ARIEL spacecraft payload module

The ARIEL mission is a European space project that aims to detect exoplanets with a spacecraft orbiting around the L2 point of the Sun-Earth system. The main payload consists of a Cassegrain telescope composed of mirrors that reflect and concentrate the incoming light from the deep space observations to finally guide it to the detectors. As in many other space missions, a dedicated complex assessment is established during the design phase to evaluate the impact on the optical performance caused by thermoelastic effects, which involves the coordinated work of the thermal, structural, and optical engineers. Despite that well-known and standardized processes and tools are established separately in each involved area, there is a lack of standardization about the way of exchanging the data between them, where additional calculations are required in some cases. This work focuses on the temperature mapping, which is the intermediate step between thermal and structural analyses, where temperatures are transferred to the structural model. The main difficulty of this process is related to the differences in modelling methods and approaches between both models, being necessary the development of an adequate algorithm to find the most accurate transfer of temperatures. This paper shows two different options for temperature mapping, detailing the proposed flowcharts. One of these methods requires the performance of an additional thermal conductive analysis, where a new improved procedure has been implemented in this work to solve some computational issues that made its application for large models difficult or even unfeasible. Both temperature mapping methods have been applied to the payload module of the ARIEL spacecraft, comparing the output results in terms of temperatures, stresses, forces, and displacements to evaluate their differences.

Modeling and control of two-stage magnetically suspended rotating payload platform

The new generation of remote sensing satellites improves the coverage of remote sensing imaging by improving the orbital surface and using the rotary scanning imaging technology. In this paper, a new technical scheme of rotating payload platform based on disturbance-free payload technology is proposed to adapt to the rotary scanning imaging satellite. The platform adopts the cubic Stewart configuration and uses the Lorentz force magnetic actuator as the actuator to form the satellite support module and the payload into a noncontact two-stage three-body. First, this paper uses Newton-Euler method to establish a six-degree-of-freedom dynamic model for the new rotating payload platform, and establishes a kinematic model by quaternion method. Then, according to the established kinematics and dynamics model, the control scheme of absolute pointing control of payload platform (upper platform) and relative follow-up control of kinetic energy container (middle platform) is designed. Furthermore, the six-degree-of-freedom decoupling control method of the actuator is designed. Based on this, the absolute pointing controller of the payload module and the relative position and attitude calculate controller of the kinetic energy container are designed. Finally, the research of simulation is carried out. The simulation results show that the new rotating payload platform has higher rotation pointing accuracy and stability than the traditional ultra-quiet platform, and the vibration isolation acting on the support module and the vibration absorption acting on the payload are more effective.

In-Flight Calibration of Lorentz Actuators for Non-Contact Close-Proximity Formation Satellites with Cooperative Control

The non-contact close-proximity formation satellite (NCCPFS) is one of the technical solutions to improve the attitude performance, consisting of a payload module (PM) and a support module (SM). The non-contact Lorentz actuator (NCLA), as the core components of the NCCPFS, directly affect the attitude control performance of the entire satellite. In order to ensure the ultra-high attitude pointing performance and stability of the PM, an in-flight calibration method for the NCLAs based on minimum model error (MME) algorithm and Kalman filtering (KF) with cooperative control strategy is proposed in this article. In this method, the NCLAs generate a periodic nominal torque that causes the attitude of the PM to be periodically deflected. This periodic torque also reacts on the SM, and the SM counteracts this periodic torque through the flywheel to realize the cooperative tracking relative to the PM. Then, the gyroscope data are substituted into the MME algorithm to obtain the angular acceleration of the two modules, and the KF algorithm is adopted to observe the actual output torque of the NCLAs to complete the in-flight calibration of the NCLAs. Numerical simulation results show that the accuracy of the proposed calibration algorithm can reach about 8%, which proves the effectiveness of the proposed method.

Passivity-based fractionated payload 6-DOF maneuver control using electromagnetic actuation

Employing fractionated spacecraft architecture with electromagnetic actuation into space missions can enhance resilience and flexibility since all subsystems of conventional monolithic spacecraft are separated as several modules to be easily replaced on-orbit when failure happens. Previous electromagnetic configuration used in fractionated spacecraft would render unexpected forces and torques which may undermine the stability of entire system. This paper considers the maneuver of a separated payload in a fractionated spacecraft architecture (FSA) using electromagnetic actuation with adaptive passivity-based control. The FSA consists of a fractionated payload module (FPM) and a supporting spacecraft module (SSM). Relative motion dynamics of FPM and SSM are formulated on SE(3). An adaptive passivity-based 6-DOF motion control scheme is established for FPM's translation and rotation tracking control simultaneously. The closed-loop equilibrium can be attained by the proposed passivity-based 6-DOF control scheme. The controller has no requirement for the precise inertia knowledge of FPM, which processes the robustness of the modelling errors. Numerical simulations are presented to show the effectiveness of the proposed control scheme.

Design and Verification of Rotating Avoiding Type Locking-Releasing Mechanism

Given the resistance environment and rotation requirements of the satellite payload module, a horseshoe-shaped compression slider bearing method and a double-layer structure shear resistance method are proposed. A rotating avoiding type locking-releasing mechanism is designed, which has a large bearing capacity and reliable locking, and can leave a sufficient rotation avoidance gap for the payload module after release. Then, the analysis and design process of preload is given, and the axial tensile and transverse shear tests are carried out to verify the reasonableness and feasibility of the designing scheme, which solved the technical problem of rotating operation of the 2-ton payload module after enduring the harsh mechanical environment effectively, and can be applied to the task of locking and releasing the same kind of rotating type payload in the resistance environment.

Formation of an approach to modeling orbital operations assembly of a reconfigurable spacecraft on geostationary orbit

The aim of the study is to form an approach to modeling the operations of the orbital assembly of a reconfigurable spacecraft (RS) in geostationary orbit. Reconfigurable spacecraft are a set of modular spacecraft (MS), where, in a particular case, one MS can be assigned the functions of the service systems module (MSS), and the second - the functions of the payload module (MPN). To ensure the assembly of the RC, or the replacement of some MC, for example, in case of its failure with a new one, it is necessary to provide a solution to the problem of bringing the MS with the RS. The article analyzes and studies the operation of the motion control system of the MS during the convergence of the MS with the RS. A list of necessary mathematical models for performing operations in solving the problem of convergence of the MS with the RS is formed, and a block diagram of the interaction of mathematical models is presented. The paper presents a brief description of the mathematical apparatus that allows modeling the operations of convergence of the MS with the RS. This mathematical apparatus includes: a model of the orbital motion of the MS and the RS, models of the angular motion of the MS and the RS, sensitive elements and executive bodies. In this paper, the mathematical modeling of the MS with the RS convergence operations is considered as the subject of research. The object of the study is the motion control system of the MS, which ensures the implementation of the approach of the RS in geostationary orbit.

Conceptual design and control method for a non-contact annular electromagnetic stabilized satellite platform

Disturbance-Free Payload (DFP) spacecraft can meet the requirements of ultra-high attitude pointing accuracy and stability for future space missions. However, as the main control actuators of DFP spacecraft, Linear Non-Contact Lorentz Actuators (LNCLAs) have control output problems with six-degree-of-freedom coupling and nonlinear effects, which will affect the attitude control performance of DFP spacecraft. To solve this problem, a novel concept for Non-Contact Annular Electromagnetic Stabilized Satellite Platform (NCAESSP) is proposed in this study. The concept is centered on replacing the LNCLAs with a non-contact annular electromagnetic actuator to solve the two problems mentioned above. Furthermore, for the different control requirements of the payload module and the support module of the NCAESSP, a high-precision attitude controller based on the robust model matching method and a dual quaternion-based adaptive sliding mode controller are proposed. Additionally, the simulation results verify the feasibility and effectiveness of the proposed approach.

Pixel source follower glow in HxRG detector

AbstractThe Atmospheric Remote‐Sensing Infrared Exoplanet Large‐survey (ARIEL) space mission is dedicated to the study of the exoplanets' atmosphere. To do so, the payload module is made of two instruments. The ARIEL InfraRed Spectrometer instruments is one them, it will perform spectroscopic measurements from the near‐infrared domain to the long‐wave infrared domain. The instrument relies on the use of two H1RG detectors. During early phases of instrument development dark current measurements have been done on an 8 m cutoff detector operated in window mode. These measurements revealed an unexpected level of dark current. This observation triggered a joint series of tests at the European Space Agency technical center and at the Astrophysics Department of the Atomic Energy Commission. As demonstrated in this paper, the excess of dark current originates from the collection, by the photosenstive layer, of light emitted by the source follower transistor of the pixel unit cell. This effect is referred as glow. Based on experimental results, this glow is studied as a function of the detector operating conditions (temperature, biases, timing diagram). These results also show that for the range of operating temperatures considered, the dark current performance of the detector is limited by the glow.

Life Cycle Assessment of the Cellulosic Jet Fuel Derived from Agriculture Residue

The purpose of this paper is to discover the impacts of contradictory factors in the application of agricultural residue with sustainable biofuel benefits. Based on the Life cycle assessment (LCA) approach, the quantitative LCA assessment model and approach have been established, coupling upstream cultivation and downstream jet biofuel product, which would benefit agriculture residue choice. The LCA model investigated the effects of interaction factors on energy consumption, including land release and agriculture residue use change. The computational framework of the LCA model is classified into three sub-models, including the cultivation and harvesting model, the refining process and distribution model, and the flight model. According to uncertainty analysis by the LCA model, the positive energy gains have been conducted at a wide range of hydrogen production and methanol production. The application model is represented by six types of typical aircraft widely used in China, including the LTO cycle module, actual cruising distance and maximum cruising distance module, actual payload, and maximum payload module. In the whole life cycle assessment, GHGs of agriculture residue is 17.9 gCO2e/MJ while petroleum-based jet fuel is 90.2 gCO2e/MJ. The order of GHGs in WTW (well to wheel) is agriculture residue < corn stover < beanstalk < wheat straw < rice straw. The land release conducted obviously to the total GHGs emission for rice straw, which indicated that land release should involve in the LCA.

Design of a thermal control system with a two-phase coolant for a spacecraft with a capacity of 7 kW

This article presents a methodological approach to designing a spacecraft thermal control system with coolant pumping with a cooling capacity of up to 7.0 kW. Two design options are considered. There is no airtight instrument container in the spacecraft layout scheme, and all heat-generating equipment is located directly on the power structure panels, so excess heat is removed from the spacecraft directly from the outside of the instrument panels. With all the attractiveness of a two-phase circuit with a heat pipe circuit, its use in automatic spacecraft is complicated by the need to supply concentrated heat to the capillary evaporator of the circuit. To do this, it is necessary to collect heat from a large surface of the structure, on which a large number of heat sources are installed. A schematic solution of the thermal control system is considered, in which the thermal power of the payload module is distributed between the panel of the service systems module and deployable radiators. At the same time, in the first version, the heat pipes of the payload module are connected to the heat pipes of the service systems module along the profile shelves, the contour heat pipes of the deployable radiator are connected only to the heat pipes of the service systems module. Ums, the heat load of the payload module is transferred to the heat pipes of the service systems module and then to the loop heat pipes. The second option differs from the first one in that to equalize the temperatures of the panels of the payload module, each heat pipe of the North panel is connected to the heat pipes of the South panel. From the results of a comparative analysis of the mass budget and energy efficiency, the conclusion follows: the DFC with capillary pumping is the most preferable option 2. which, with the same other characteristics, has a smaller mass. The specific mass-energy: characteristic of such a system is ~22.9 kg/kW.

An Extended Configuration Design of High-Throughput Satellite

The high-throughput satellite (HTS) adopts multi-spot beam technology, which enables the satellite design process to be featured by a large number of antenna horns, large-scale arrays, heavy support structures, and complex interfaces with the transponder. To solve the layout problem of complex payloads of high-throughput satellites, this paper proposes a configuration design scheme for the satellite expansion module. With the configuration design optimization of the high-throughput satellite payload module of the DFH-4E platform serving as an example, the mechanical analysis software PATRAN was used for modeling and mechanical analysis, and the configuration was optimized according to the analysis results. The optimization results show that the configuration design scheme of the expansion module is adopted to solve the layout requirements of high-throughput payloads effectively. The relevant indicators will meet the requirements of the system. And the design scheme is reasonable and feasible, which can provide a reference for the configuration design of high-throughput satellites.

Probing cosmic inflation with theLiteBIRDcosmic microwave background polarization survey

AbstractLiteBIRD, the Lite (Light) satellite for the study of B-mode polarization and Inflation from cosmic background Radiation Detection, is a space mission for primordial cosmology and fundamental physics. The Japan Aerospace Exploration Agency (JAXA) selected LiteBIRD in May 2019 as a strategic large-class (L-class) mission, with an expected launch in the late 2020s using JAXA’s H3 rocket. LiteBIRD is planned to orbit the Sun–Earth Lagrangian point L2, where it will map the cosmic microwave background polarization over the entire sky for three years, with three telescopes in 15 frequency bands between 34 and 448 GHz, to achieve an unprecedented total sensitivity of $2.2\, \mu$K-arcmin, with a typical angular resolution of 0.5○ at 100 GHz. The primary scientific objective of LiteBIRD is to search for the signal from cosmic inflation, either making a discovery or ruling out well-motivated inflationary models. The measurements of LiteBIRD will also provide us with insight into the quantum nature of gravity and other new physics beyond the standard models of particle physics and cosmology. We provide an overview of the LiteBIRD project, including scientific objectives, mission and system requirements, operation concept, spacecraft and payload module design, expected scientific outcomes, potential design extensions, and synergies with other projects.

Integrated Control Scheme for an Improved Disturbance-Free Payload Spacecraft

For a novel disturbance-free payload (DFP) spacecraft, it is difficult to isolate the low-frequency disturbances owing to the umbilical cables, which decreases the pointing accuracy and stability of the payload. In this research, an improved DFP spacecraft and its integrated control scheme are designed to enhance the pointing accuracy and disturbance attenuation performance. The improved DFP spacecraft consists of a Payload Module (PM), a Support Module (SM), and a Test Mass (TM). The integrated control system is subdivided into three interconnected control loops. An active vibration isolation control loop is used to isolate the PM from disturbances in the high-frequency bands and control the PM to track the attitude of the SM. A drag-free control loop is used to isolate the SM from disturbances in the low-frequency bands and control the SM to track the attitude of the TM. An attitude-pointing control loop is used to control the TM to track the desired attitude. Based on the improved DFP spacecraft and the integrated control system, the payload mounted on the PM can be isolated from disturbances in all of the frequency bands, and its high-level requirements for pointing accuracy and stability can be realized.

Homogeneity of antibody-drug conjugates critically impacts the therapeutic efficacy in brain tumors.

SUMMARYGlioblastoma multiforme (GBM) is the most aggressive and fatal disease of all brain tumor types. Most therapies rarely provide clinically meaningful outcomes in the treatment of GBM. Although antibody-drug conjugates (ADCs) are promising anticancer drugs, no ADCs have been clinically successful for GBM, primarily because of poor blood-brain barrier (BBB) penetration. Here, we report that ADC homogeneity and payload loading rate are critical parameters contributing to this discrepancy. Although both homogeneous and heterogeneous conjugates exhibit comparable in vitro potency and pharmacokinetic profiles, the former shows enhanced payload delivery to brain tumors. Our homogeneous ADCs provide improved antitumor effects and survival benefits in orthotopic brain tumor models. We also demonstrate that overly drug-loaded species in heterogeneous conjugates are particularly poor at crossing the BBB, leading to deteriorated overall brain tumor targeting. Our findings indicate the importance of homogeneous conjugation with optimal payload loading in generating effective ADCs for intractable brain tumors.

Event-Triggered Attitude-Tracking Control for a Cableless Non-Contact Close-Proximity Formation Satellite

The cablelessness of non-contact close-proximity formation satellites can fundamentally avoid the influence of non-contact interface coupling effects and can further enhance the attitude pointing accuracy and stability of the payload module (PM). However, it also brings the problem of limited on-board resources and system latency. In this paper, an event-triggered attitude tracking controller of the support module (SM) that avoids the Zeno phenomenon was proposed. The update time of the control signal was determined by the event-triggering mechanism based on intermediate variables, thus, reducing the communication burden and actuator asynchrony between the two modules. The feasibility and effectiveness of the proposed approach was demonstrated by numerical simulations.

Stringent three-axis stability control for novel non-contact satellite using embedded second-order noise estimator

To achieve the stringent three-axis stability performance, a non-contact satellite framework composed of the payload module and bus module is designed and an embedded second-order noise estimator based controller is proposed in this paper. For the payload module, the embedded second-order noise estimator is developed to envelope the nonlinear three-axis coupling terms, and the stringent attitude control is implemented by adopting the non-contact Lorentz actuators. For the bus module, a collision avoidance controller is adopted to maintain the relative motion between the payload module and bus module. The hardware experiments based on air-bearing platform have been conducted to verify the feasibility and effectiveness of the proposed approach, and the results show that the new methods have superior and reliable performance with low computational cost.

Compound Attitude Maneuver and Collision Avoiding Control for a Novel Noncontact Close-Proximity Formation Satellite Architecture

In order to achieve the attitude maneuver performance of the noncontact close-proximity formation satellite architecture, this paper presents a compound control strategy with variable-parameter sliding mode control and disturbance observer-based feedforward compensation. Firstly, the variable-parameter sliding mode control is proposed to guarantee the attitude maneuver performance of the payload module. Secondly, the collision avoiding control with disturbance observer-based feedforward compensation is proposed to guarantee the synchronization of the two separated modules within the small air clearance constraint of the noncontact Lorentz actuator. Finally, a physical air-floating platform is established to verify the effectiveness of the proposed approach.

Dynamic Modeling and Control Method of Dual-Body Satellite

For the attitude control problem of dual-body satellite with non-solid connection between payload module and platform module, this paper proposes a dual-body dynamics modeling and control method based on the concept of disturbance-free payload(DFP). The payload module is rotated in orbit with high precision, and its forces can be determined by performing a force analysis on it. The Euler equation is used to calculate the attitude information of the payload module, so that the motion of the payload module can be described more accurately. By designing the control algorithm and the control loop with the idea of DFP, the attitude accuracy of the payload module is better controlled. The simulation results show that the controller can keep the attitude accuracy of the payload module within “1.65”, the results meet the expectation.

Three-dimensional inhomogeneous thermal fields of the “Photon-Amur 2.0” payload electronic board developed for nanosatellites

The thermal fields of the Photon-Amur 2.0 payload electronic board developed for nanosatellites were studied. The Photon-Amur 2.0 payload consists of an electronic control board with a casing mounted in a nanosatellite and a remote panel with experimental photovoltaic converters. A modified heat balance method was used for numerical simulation of the thermal fields of the control board and the casing. The constructed model and the obtained results of the numerical simulation were verified by comparison with the thermal diagrams obtained for the Photon-Amur 2.0 electronic board under normal operating conditions. For modeling the outer space operating conditions, it was assumed that there is a vacuum outside and inside the Photon-Amur 2.0 casing, and the thermal effect is transmitted from the nanosatellite racks to the payload electronic board through the fastenings. The thermal effect is of a periodic nature with amplitude of 45 to +80○C and a period of 96 min, which approximately corresponds to the motion of a nanosatellite in a 575 km-high orbit. It was demonstrated that with such composition of the payload module, its casing can work as a passive thermoregulator of thermal fields on the electronic board of Photon-Amur 2.0. The simulation showed that the casing helps to keep the temperature on the control board in the interval of 15C to +85C, which is acceptable for the electronic components used on the payload control board.

Concept of Short-Range Radar and Counter-Projectile for Detecting and Countering Armour-Piercing Projectiles

During warfare and acts of terrorism an extreme threat to vehicles and other high-value assets comes from armour-piercing projectiles. Under these conditions, defence systems should include devices capable of rapid detection of these threats. Defence assets should also be provided with counter-projectile systems capable of destroying incoming armour-piercing projectiles at a safe distance from the asset to be protected. This paper describes the concept of a system comprising of a lightweight short-range radar and a counter-projectile for countering armour-piercing projectiles. The purpose of the radar is to monitor the environment and search for incoming armour-piercing projectiles. When an armour-piercing projectile is detected in a designated monitoring area, an automatic command is given for the counter-projectile launcher to be fired. The counter-projectile deployed can be equipped with a single or multi-sensor detection head unit and an explosive payload module, both being the primary components of the warhead. When the signal analysis blocks interfaced with the detection head determine that the armour-piercing projectile to be struck down is in the target position in relation to the counter-projectile deployed, they automatically command the explosive payload module to detonate. The components of the system concept were tested in proving ground conditions. The successful results of these tests confirmed the validity of the solutions initially adopted and the execution of the individual systems.